Hybrid propulsion system

ABSTRACT

Disclosed is a propulsion system having a structural configuration that provides easy and convenient access to the interior regions of a liquid fuel tank and a hybrid rocket motor case. In one embodiment, the propulsion system comprises: a hybrid rocket motor case and a fuel tank coupled to the hybrid motor case. The motor case is configured to hold solid rocket fuel and the fuel tank defines an internal volume configured to hold a fluid oxidizer. A bulkhead is interposed between the motor case and the fuel tank, wherein at least one access passageway extends through the bulkhead. The access passageway provides exterior access to the interior volume of the motor case or the internal volume of the storage tank while the hybrid rocket motor is coupled to the fuel tank.

RELATED APPLICATION

This application claims the benefit of priority to U.S. ProvisionalPatent Application Ser. No. 60/614,789, entitled “Hybrid PropulsionSystem”, filed Sep. 29, 2004, which is incorporated herein by referencein its entirety.

TECHNICAL FIELD

This disclosure relates to propulsion systems, and more particularly toa hybrid propulsion system.

BACKGROUND

Liquid systems and solid systems are the two basic types of rocketpropulsion systems that are generally used in the rocket industry. In asolid propellant system, solid rocket fuel and an oxidizer are mixedtogether and allowed to cure inside a rocket case to form a solidpropellant material, which is then ignited in the rocket case. Uponignition, pressure forms within the rocket case and gases are releasedthrough a nozzle to produce thrust. In a solid propellant system, thesolid propellant burns uninterrupted until all the propellant isexhausted, which can be undesirable in certain circumstances. Solidsystems can be quite complicated, and are subject to severalrequirements during manufacture in order to minimize safety risks duringuse. For example, the solid propellant must be crack-free, as propellantgrains which contain cracks present a risk of explosive failure of thevehicle. Solid propellant systems can be inadvertently ignited bymechanical shock and static electricity. Consequently, the manufacturingprocess requires extreme safety precautions, which increasesmanufacturing costs.

In a liquid system, a liquid oxidizer is fed into a combustion chamberin combination with a liquid fuel. The oxidizer and liquid fuel aremixed in the combustion chamber, where they react to produce gases underhigh temperature and high pressure. The gases exhaust through a nozzlefrom the combustion chamber to thereby produce thrust. Although widelyused, there are certain drawbacks associated with liquid propulsionsystems. One such drawback is that the mixing of reactants requires ahigh performance pressurization system for the fuel and oxidizer, whichcan contributes to a high cost with respect to both money andmaintenance. Like solids, a liquid system can also explode since theoxidizer and fuels can be inadvertently mixed together. Another drawbackis that exotic—and therefore expensive—materials must be used for thevarious components of the system, which increases the monetary cost ofthe systems.

Another type of rocket propulsion system are the hybrid systems, whichare generally not as widely used as liquid and solid rocket fuelsystems. A hybrid system combines aspects of both liquid systems andsolid systems in that one propellant is stored as a solid and anotherpropellant is stored as a liquid. In a typical system, the solidmaterial is used as the fuel and the liquid material is used as theoxidizer. A variety of materials can be used as the fuel, includingPlexiglas (polymethyl methacrylate (PMMA)), high density polyethylene(HDPE), hydroxyl terminated polybutadiene (HTPB), etc. Nitrous Oxide isa commonly used as the oxidizer, although other oxidizers can be used.

Hybrid systems have characteristics that can be highly desirable forcertain situations and uses. For example, a hybrid system generally hashigher specific impulse than solid systems. Specific impulse is thechange in momentum per unit mass for the rocket fuel. Thus, a hybridsystem can generate a high level of “push” for each unit of fuel that isused. Another advantage associated with hybrid propulsion systems is thecomplete separability of the fuel from the principal oxidizer. Thisinhibits the potential for inadvertent ignition or catastrophic failureso that hybrid systems are inherently immune to inadvertent explosion.Yet another advantage is that hybrid systems have the ability to easilystart, stop, and restart the combustion of the rocket fuel.

There is currently a need for improved hybrid propulsion systems.

SUMMARY

Disclosed is a propulsion system having a structural configuration thatprovides easy and convenient access to the interior regions of a liquidfuel tank and a hybrid rocket motor case. In one aspect, the propulsionsystem comprises: a hybrid rocket motor case having a proximal end and adistal end, the motor case defining an interior volume including acombustion chamber at the proximal end, wherein the motor case isconfigured to hold solid rocket fuel; a fuel tank coupled to theproximal end of the motor case, the fuel tank defining an internalvolume configured to hold a fluid oxidizer; and a bulkhead interposedbetween the motor case and the fuel tank, wherein at least one accesspassageway extends through the bulkhead, the access passageway providingexterior access to the interior volume of the motor case or the internalvolume of the storage tank while the hybrid rocket motor is coupled tothe fuel tank.

In another aspect, the propulsion system comprises: a hybrid rocketmotor case having a proximal end and a distal end, the motor casedefining an interior volume including a combustion chamber at theproximal end, wherein the motor case is configured to hold solid rocketfuel; a fuel tank coupled to the proximal end of the motor case, thefuel tank defining an internal volume configured to hold a fluidoxidizer; and a bulkhead interposed between the motor case and the fueltank, the bulkhead including means for providing exterior access to theinterior volume of the motor case or the internal volume of the storagetank while the hybrid rocket motor is coupled to the fuel tank.

The details of one or more embodiments are set forth in the accompanyingdrawings and the description below. Other features, objects, andadvantages of the invention will be apparent from the description anddrawings, and from the claims.

DESCRIPTION OF DRAWINGS

FIG. 1 is a schematic view of a single port hybrid rocket motor.

FIG. 2 is a schematic view of a first embodiment of a hybrid propulsionsystem.

FIG. 3 is an enlarged view of a bulkhead boundary between a combustionchamber casing and an oxidizer storage tank of the system of FIG. 2.

FIG. 4 is a schematic view of a nozzle of the system showing a liquidinjection thrust vector control configuration

FIG. 5 is a schematic view of a hybrid propulsion system that employs acold gas attitude and control system.

FIG. 6 shows a schematic view of a first embodiment of a cold gasattitude and control system.

FIG. 7 shows a schematic view of a first embodiment of a cold gasattitude and control system.

FIG. 8 is a perspective view of the MTV.

FIG. 9 is a perspective view of the MTV coupled to a payload.

DETAILED DESCRIPTION

FIG. 1 shows a schematic of the configuration of a single port hybridrocket motor 100. The motor 100 generally includes a main casing 102that encloses a main chamber having a combustion chamber 110, a mixingchamber 115, and an elongate combustion port 120 extending therebetween.A solid fuel 130 is located within the main chamber. An injector 122communicates with the combustion chamber 110 for injecting a liquidphase oxidizer into the combustion chamber 110. In use, the oxidizer isinjected into the combustion chamber 110 via the injector 122. Theinjected oxidizer is gasified and flows axially along the combustionport 120, forming a boundary layer edge 125 over the solid fuel 130. Theboundary layer edge 125 is usually turbulent in nature over a largeportion of the length of the combustion port 120. A diffusion flame zone135 exists within the boundary layer edge 125, which diffusion flamezone 135 extends over the entire length of the solid fuel 130.

The heat generated in the flame, which is located approximately 20-30%of the boundary layer thickness above the fuel surface, is transferredto the wall mainly by convection. The wall heat flux evaporates thesolid fuel and the resultant fuel vapor is transported to the flamewhere it reacts with the oxidizer, which is transported from the freestream by turbulent diffusion mechanisms. The unburned fuel that travelsbeneath the flame, the unburned oxidizer in the free stream, and theflame combustion products mix and further react in the mixing chamber115. The hot gases expand through a nozzle 140 to deliver the requiredthrust.

A hybrid rocket motor 100 of the type shown in FIG. 1 can beincorporated into various types of propulsion systems. For example, FIG.2 shows a hybrid propulsion system 202 having an elongate motor 100attached to at least one storage tank 215 that holds a fluid oxidizer ina liquid phase. The motor 100 is at least partially configured asdescribed above with respect to FIG. 1. In the illustrated embodiment,the casing 102 of the motor is attached at a first end to the storagetank 215 such that the nozzle 140 is spaced from the storage tank 215.In this manner, the motor extends outwardly from the storage tank 215.It should be appreciated that the shape of the storage tank can vary.

An intermediate structure, such as a bulkhead 211, is interposed betweenthe casing 102 and the storage tank 215 and provides access to theinterior of the casing 102 and the interior of the storage tank 215, asdescribed more fully below. Advantageously, neither the casing nor thestorage tank 215 has to be removed to gain access to the interiorregions of the components. Rather, access to the interior of the storagetank 215 and the casing 102 is provided via passageways that extendthrough the bulkhead 211, as described more fully below. The bulkhead211 can be attached to the storage tank 215 and the casing 102 in avariety of manners, such as, for example, bolts, screws, adhesive,clamps, etc., or combinations thereof.

FIG. 3 shows an enlarged, side view of the region of attachment betweenthe storage tank 215 and the casing 102 contained in circle 302 in FIG.2. With reference to FIG. 3, the bulkhead 211 is interposed between thecasing 102 and the motor 100. The bulkhead 211 is substantially circularwhen viewed from the front. The bulkhead 211 has a circumferentialflange 213 that extends around the entire circumference of the bulkhead211 so as to form a region that can be attached to the casing 102 andthe storage tank 215. Specifically, the flange 213 functions as aretaining region that can be used to attach the wall of the tank 215 tothe wall of the casing 102 with the bulkhead 211 interposedtherebetween. It should be appreciated that the bulkhead 211 sealinglyengages the casing 102 and the storage tank 215. This prevents gases orany material from escaping from either the storage tank or the casing atany location along the attachment region between the bulkhead and thestorage tank/casing.

As mentioned, a retaining means, such as bolts, screws, glue, etc., orcombinations thereof, can be used to attach the tank 215 and the casing102 to the flange 213 of the bulkhead 211. At least a portion of thebulkhead 211 has an outwardly-facing region, such as an outer wall 221,that can be accessed from the exterior of the rocket motor 100 and fromthe exterior of the storage tank 215. In the illustrated embodiment, theouter wall 221 comprises an outwardly-facing side wall of the flange213. Thus, an operator or technician can access the outer wall 221without having to remove either the storage tank 215 or the casing 102from the bulkhead 211.

With reference still to FIG. 3, one or more access passageways extendthrough the bulkhead 211 to provide access to the interior of thestorage tank 211 and/or the interior of the rocket casing 102 from alocation exterior to the rocket motor. The access passageways extendthrough the bulkhead 211 to provide access to the interior of the casing102 and/or to the interior of the storage tank 215 via theoutwardly-facing sidewall 221 of the bulkhead 211.

In one embodiment, the passageways include at least one storage tankaccess passageway 223, which has at least a first opening in the outerwall 221 of the bulkhead 211 and at least a second opening thatcommunicates with the interior of the storage tank 215. Theconfiguration of the storage tank access passageway can vary. Forexample, the storage tank access passageway 223 can be a singlepassageway with no branches such that it has only one opening in thesidewall 221 and only one opening into the interior of the storage tank215. Alternately, the storage tank access passageway 223 can havebranches such that it has one or more openings in the side wall 221and/or one or more openings into the storage tank 215. In this regard,the passageway can have one or more “forks” along its length.

In addition, the storage tank access passageway does not have to extendentirely through the bulkhead 211 along its entire length. Rather, atleast a portion of the storage tank access passageway 223 can extendthrough the wall of the storage tank 215, as long the storage tankaccess passageway provides access to the interior of the storage tank215 via the bulkhead 211. Any quantity of storage tank accesspassageways can be used.

With reference still to FIG. 3, the bulkhead 211 can also have at leastone combustion chamber access passageway 229 which has at least a firstopening in the outer wall 221 and at least a second opening thatcommunicates with the interior of the casing 102. As discussed abovewith respect to the storage chamber access passageways 223, theconfiguration of the combustion chamber access passageway 229 can vary.For example, the combustion chamber access passageway 229 can be asingle passageway with no branches such that it has only one opening inthe sidewall 221 and only one opening into the interior of thecombustion chamber 110. Alternately, the combustion chamber accesspassageway 229 can have branches such that it has one or more openingsin the side wall 221 and one or more openings into the combustionchamber 110. In addition, the passageway does not have to extendentirely through the bulkhead 211 along its entire length, but can alsoextend through the wall of the casing 102 to provide access to thecombustion chamber. Any quantity of combustion chamber accesspassageways can be used.

The access passageways 223, 229 provide convenient access to theinterior of the storage tank 215 and to the combustion chamber 110. Anelongate structure, such as a tubing or catheter, can be insertedthrough either of the passageways in a permanent or temporary manner,wherein the elongate structure can be equipped with sensors and/orvisual aids that can be used to gather data regarding the interior ofthe storage tank and/or the combustion chamber. Advantageously, thestorage tank 215 and the casing 102 do not have to be disattached fromone another in order to gather such data, as the data is gathered viathe appropriate access passageways through the bulkhead 211. Thepassageways can be plugged when not in use or can be equipped withvalves or doors. In this manner, the storage tank 215 does not have tobe removed from the bulkhead 211 or the rocket tube casing 102 in orderto gain access to the interior of the storage tank 215. Likewise, thecasing 102 does not have to be removed in order to gain access to thecombustion chamber 110 in the casing interior. Rather, the passageways223 and 229 provide such access through the bulkhead 211.

With further reference to FIG. 3, the bulkhead 211 forms a boundarybetween the combustion chamber 110 and the interior of the storage tank215. An injection mechanism 231 is disposed within the storage tank 215distal of the bulkhead 211. (The distal direction is represented by thearrow D in FIG. 3.) The injection mechanism includes an injector 233that is configured to inject the oxidizer from the storage tank 215 intothe combustion chamber 110 toward the solid fuel 130 (shown in FIG. 1).The injector 233 is coupled to a valve 234 that regulates the flow ofoxidizer into the injector 233. An inlet is coupled to the valve 234 fordirecting oxidizer into the valve 234 and thereafter into the injector233. It should be appreciated that the entire injection mechanism 231 ispositioned within the storage tank 215 such that it is entirelysubmerged within the oxidizer that is located in the storage tank 215.In addition, the injection mechanism 231 is positioned distal of thecombustion chamber 110 such that the injection mechanism 231 is outsidethe boundaries of the combustion chamber 110. This differs from aconventional injector (such as the injector 122 shown in FIG. 1), wherethe injector is at least partially positioned within the combustionchamber.

With reference still to FIG. 3, when the valve 234 is opened, the liquidoxidizer in the storage tank 215 flows into and through the injector 233via the valve 234. The injector 233 injects the liquid oxidizer into thecombustion chamber 110. As discussed above, the injected oxidizer isthen gasified and flows axially along the combustion port of thecombustion chamber 110. Resultant hot gases expand through the nozzle140 to deliver the required thrust.

During the combustion process, the combustion chamber 110 achievesfluctuations or oscillations in pressure as the liquid oxidizer isgassified. In certain circumstances, it can be generally desirable tomaintain the pressure oscillations at or below a predeterminedpercentage of a reference, such as lower than about 5 percent pressureoscillations. The reason for this is that increased pressureoscillations cause vibrations, which can lead to structural failure ifthe pressure oscillations increase to too high of a level. Applicant hasobserved that it can be advantageous to intentionally increase thepressure oscillations within the combustion chamber in order to achieveincreased combustion instability within the combustion chamber 110. Anincrease in specific impulse of the rocket motor results when thepressure oscillations are increased to greater than about 5% within thecombustion chamber 110. The increase in pressure oscillations can beaccomplished by increasing the oxidizer flux, which is defined as theoxidizer flow rate divided by the cross-sectional area of the combustionchamber. Thus, it can be desirable in certain circumstances to increasethe percentage of pressure oscillations in the combustion chamber andthereby increase combustion instability, which results in a greaterspecific impulse for the rocket motor.

As discussed above, the rocket motor produces thrust as hot gases expandin a plume through the nozzle 140. With reference to FIG. 4, a resultantthrust vector TV is achieved relative to the nozzle 140. At least aportion of the liquid oxidizer from the storage tank 215 can be routedvia a conduit into the nozzle 140 to provide control over the directionof the thrust vector relative to the nozzle 140. This is described inmore detail with reference to FIG. 4. A conduit 402 provides a flow pathfor liquid oxidizer to flow from the storage tank 115 into the nozzle140. In this regard, the conduit 402 can comprise a pathway or a seriesof interconnected pathways that extend from the storage tank 115 to thenozzle 140. It should be appreciated that one or more valves can belocated along the conduit 402 so that an operator can selectively flowoxidizer from the storage tank 215 into the nozzle 140. In addition,more than one conduit 402 can be used to provide oxidizer flow intodifferent regions of the nozzle 140.

During expansion of the hot gases through the nozzle 140, liquidoxidizer is flowed into a desired location in the nozzle through theconduit 140, as represented by the arrow 404 in FIG. 4. As the liquidoxidizer flows into the nozzle 140, it combusts within the hot gases.The combustion of the liquid oxidizer results in an increase inpressure, thereby forming a high pressure region within the nozzle 140where the liquid oxidizer combusts. It should be appreciated that in theabsence of the liquid oxidizer being injected into the nozzle 140, theresultant thrust vector will point in a predetermined direction, such asis represented by the thrust vector arrow TV in FIG. 4. However, thepresence of the high pressure region resulting from the injection of theliquid oxidizer changes the direction of the resultant thrust vector,for example to the thrust vector TV1 in FIG. 4. In this manner, theliquid oxidizer can be used to vary the direction of the thrust vectorrelative to the nozzle 140. It should be appreciated that additionalconduits 402 can be located to have outlets at various predeterminedlocations in the nozzle 140 so that liquid oxidizer can be selectivelyinjected into the nozzle 140 to selectively vary the direction of thethrust vector.

Thus, there has been described at least two uses for the liquid oxidizerin the hybrid propulsion system. The first use is for combustion withinthe combustion chamber 110, which results in thrust through the nozzle140. The second use is to provide control over the direction of thethrust vector by injecting liquid oxidizer into the nozzle 140. There isnow described two additional uses for the liquid oxidizer, includingusing the liquid oxidizer for fine attitude control and using theproperties of the liquid oxidizer to drive a pump that pumps the liquidoxidizer into the combustion chamber 110.

The use of the liquid oxidizer for fine attitude control is nowdescribed. This can be used, for example, to provide fine attitudecontrol of a spacecraft that is coupled to the hybrid propulsion system.Such fine attitude control is described with reference to FIG. 5, whichshows a schematic view of a hybrid propulsion module (HPM) 210 versionof a hybrid propulsion system. It should be appreciated that theattitude control feature described herein can be incorporated into otherembodiments of hybrid propulsion systems.

With reference to FIG. 5, the HPM 210 includes one or more oxidizertanks 215 and a casing or tube 220 containing solid rocket fuel.Although the schematic in FIG. 5 shows four tanks 215, it should beappreciated that any quantity of tanks can be used. The tube 220 isconfigured according to the hybrid rocket motor configuration describedabove with respect to FIG. 1. That is, the tube 220 has an injector atone end (which communicates with the oxidizer tanks 215) and a nozzle atan opposed end. The tube 220 can be coupled to the storage tank 215 inthe manner described above in FIG. 2 or it can be coupled in a differentmanner. Oxidizer in a liquid phase from the tanks 215 enters the tube220 through the injector and combustion products emerge from the tankthrough the nozzle for generating thrust, as will be known to thoseskilled in the art. At least one igniter (not shown) is coupled to thetube 220 for igniting the liquid fuel within a combustion chamber of thetube 220, as described above.

An aft manifold system 225 (comprised of one or more conduits throughwhich the oxidizer can flow) couples each of the oxidizer tanks 215 to amotor conduit 217. The manifold system 225 and the motor conduit 217collectively provide at least one pathway through which oxidizer cantravel into the tube 220 from the tanks 215. As mentioned, an injector(not shown) is positioned at the entryway to the tube 220 in between themotor conduit 217 and the tube 220 for injecting oxidizer from theoxidizer tanks 215 into the tube 220. As mentioned, the tube 220 isconfigured as described above with reference to FIG. 1.

A pressure transducer 235 is coupled to the motor conduit 217. Inaddition, a main feed solenoid valve 227 is positioned along the motorconduit 217 and provides a means to control the flow of oxidizer fromone or more of the tanks 215 to the tube 220. An access device 230comprising a valve is included within or coupled to the aft manifoldsystem 225 to provide access to the oxidizer tanks 215 for filling ordraining the tanks.

With reference still to FIG. 5, a fore manifold system 240 is alsocoupled to each of the oxidizer tanks 215. The fore manifold system 240is comprised of one or more interconnected conduits through which fluidoxidizer can flow from the tanks 215. As used herein, the term “conduit”means any pathway or lumen through which fluid can flow and includespipes, tubes, etc. that can be made of any of a plurality of suitablematerials. The fore manifold system includes a pressure transducer 242,a burst disk 245, and a relief valve 250. As described below, the foremanifold system provides a pathway through which oxidizer from one ormore of the tanks 215 can flow to an attitude control system of the HPM210.

In one embodiment, the oxidizer tanks 215 house a liquid oxidizercomprised of Nitrous Oxide (N₂O). Those skilled in the art willappreciate that Nitrous Oxide is self-pressurizing at room temperature.Accordingly, the high vapor pressure of the Nitrous Oxide can beutilized in the oxidizer tanks to transport the Nitrous Oxide to thetube 220 via the aft manifold 225 and motor conduit 217 without the useof pumps or a pressurization system. Those skilled in the art willappreciate that other types of liquid oxidizers can also be used. Theoxidizer can be stored in the tanks 215 within a temperature range suchthat it exists simultaneously in both a liquid phase and a gas phase. Inthe case of Nitrous Oxide being used as an oxidizer, the Nitrous Oxideis stored in the tanks 215 at a temperature range of approximately 0degF. to 80 degF. and at a pressure of approximately 280 psia to 865psia. Such a range of pressures and temperatures is sufficient tomaintain the Nitrous Oxide within the tanks in both a liquid phase and agas phase. It should be appreciated that the pressure and temperatureranges can vary based on the substance in the tanks.

In one embodiment, the solid fuel comprises polymethylmethacrylate,although other materials can be used as the solid fuel.

As mentioned above, the hybrid propulsion module 210 further includes anattitude control system (ACS) 255, which is schematically represented bya phantom box in FIG. 5 and described in more detail below withreference to FIG. 6. In one embodiment, an ACS conduit 260 is coupled tothe aft manifold system 225. Alternately the ACS conduit may be coupledto the fore manifold system 240. Thus, the ACS conduit 260 and the aftmanifold system 225 collectively provide a pathway for the oxidizer toflow to the ACS 255 from one or more of the oxidizer tanks 215. The ACSconduit 260 provides a pathway for liquid oxidizer to flow to the ACS255 from one or more of the oxidizer tanks 215. In this regard, the ACSconduit 260 is fluidly coupled to the fore manifold system 240.

FIG. 6 shows an enlarged, schematic view of a first embodiment of theACS 255 (the remainder of the HPM 210 is represented by a phantom box210 in FIG. 5). As discussed, the ACS conduit 260 provides a pathway forliquid fuel to flow from the one or more of the oxidizer tanks 215 intothe ACS 255. A pressure regulator 310 is located along the ACS conduit260, such as at the entryway into the ACS 255. The pressure regulator310 is configured to regulate the downstream pressure of liquid fuelflowing through the main ACS conduit 260, as described in more detailbelow. The “downstream” direction is the direction toward the ACSthrusters from the tanks 215 and is represented by the arrow 312 in FIG.6. The term downstream is also used to denote relative location. Forexample, a first item or state that is located “downstream” of a seconditem or state is located in the downstream direction relative to thesecond item or state. The term “upstream” is the opposite of downstream.

An accumulator 315 can be located along the ACS conduit 260 downstreamof the pressure regulator 310 such that the oxidizer (such as NitrousOxide) is regulated by the pressure regulator 310 into the accumulator315. The accumulator 315 functions to reduce or eliminate pressurefluctuations in gas supplied to the thrusters downstream of theaccumulator.

At a location 320, the ACS conduit 260 branches into at least onethruster system 325 located downstream of the pressure regulator. Eachthruster system 325 includes one or more thrusters 330, wherein thethrusters are fed by the accumulator 315. In the illustrated embodiment,each thruster system 325 includes three thrusters 330 comprised ofsolenoid thrusters that are arranged in a triad configuration, which isdescribed in more detail below. However, it should be appreciated thatthe quantity of thrusters 330 per thruster system 325 and the number ofthruster systems 325 can vary. For example, each thruster system 325 caninclude a single thruster 330, two thrusters 330, or three or morethrusters 330.

For the first embodiment, a thrust level of approximately 0.5 lbf isenvisioned, which can be tailored by adjusting regulator set pressure.The thruster is envisioned as an on-off solenoid valve closely coupledwith a nozzle of appropriate size and expansion ratio. It should beappreciated that the thrust level and the configuration of the thrusterscan vary.

In one embodiment, the thrusters 330 do not produce thrust throughcombustion, but rather produce thrust through the expansion of cold gasexpelled from the thrusters. In this system, the cold gas is obtainedfrom the tanks 215 such that the same material that is used as theoxidizer for the rocket motor is also used as the cold gas for achievingthrust in the ACS 255. A cold gas propulsion system is desirable forfine attitude control, as such a system can provide a small minimumimpulse bit. In addition, such a system is highly reliable and safe inoperation.

As mentioned, the tanks 215 can store the oxidizer simultaneously inboth a liquid phase and a gas phase, such as in the case of NitrousOxide being used as the oxidizer. In this regard, it is desirable thatall of the oxidizer reach the thrusters 330 solely in a gas phase andthat none of the oxidizer is in the liquid phase during expulsion fromthe thrusters 330. It is generally undesirable for the thrusters 330 toexpel oxidizer in a liquid form. In other words, the oxidizer should bein a gas phase upstream of the thrusters 330 such that the oxidizerenters the thrusters 330 in the gas phase. Toward this end, the pressureregulator 310 has a set point that is below the pressure that theoxidizer can exist in a liquid state for a given temperature, whereinthe given temperature is the temperature of the oxidizer at a locationupstream of the thrusters and downstream of the pressure regulator 310,such as the temperature in the accumulator. This ensures that theoxidizer is in a gas phase and will not be in a liquid phase upstream ofthe thrusters.

In one embodiment, the temperature of the oxidizer contained in theaccumulator determines the regulator set point pressure. The pressureregulator is set for a pressure that is lower than the vapor pressure ofthe particular oxidizer at the lowest temperature of the accumulator.Thus, the pressure regulator ensures that the oxidizer will be in a gasphase in the accumulator by regulating the oxidizer pressure to apressure that is below the vapor pressure of the oxidizer for the lowesttemperature in the accumulator. Pressure regulation of the oxidizerflowing through the ACS conduit to a pressure below the oxidizer's vaporpressure at the accumulator temperature ensures that liquid will not beejected from the thrusters 330. It should be appreciated that thepressure to which the regulator 310 regulates the pressure can varybased upon the oxidizer that is being used in order to maintain theparticular oxidizer in the gas phase upstream of the thrusters 330. Toincrease the quality of the vapor downstream of the regulator, heat maybe added rather than relying on ambient heat capacity of thesurroundings.

FIG. 7 shows a simplified version of the ACS system, wherein a singleoxidizer tank 410 is coupled to a single thruster 420 via an ACS conduit260. A pressure transducer 430, pressure regulator 435, and anaccumulator 440 are located in series along the ACS conduit 260. Theconfiguration shown in FIG. 7 is similar to that shown and describedwith respect to FIG. 6 although the number of components has beenreduced.

As mentioned, the HPM 210 can be incorporated into a maneuvering andtransport vehicle (MTV) that can be used, for example, to transport apayload from a drop-off orbit to an operational orbit in space. FIG. 8shows a perspective, partial cutaway view of one embodiment of an MTV510 that incorporates the HPM 210. The MTV 510 is generally cylindrical,cubic or hexagonal in shape and includes a central motor 515 that isaligned along a longitudinal axis 520. The motor 515 is configuredsimilarly to the motor shown and described above with reference toFIG. 1. A plurality of oxidizer tanks 525 are arranged in an annularconfiguration around the central motor 515 and the longitudinal axis520. In one embodiment, there are four tanks 525 disposed around themotor 515. However, it should be appreciated that any quantity of tanks525 can be employed. A flow control valve 527 is coupled to an aftmanifold system 530 that couples the oxidizer tanks 525 to the motor515. A plurality of solar array panels 535 can be located on the MTV510. For clarity of illustration, the solar panels are not shown in FIG.8.

With reference still to FIG. 8, a fore manifold system 710 is located onan upper region of the MTV 525 above the tanks. The fore manifold system710 includes a plurality of conduits that provide pathways for liquidfuel to flow out of the tanks 525. In one embodiment, at least one gasthruster 447 is coupled to the fore manifold system via another conduit(not shown). The at least one gas thruster 447 is part of an ACS systemof the MTV 510, wherein the ACS system is configured according to thesystem described above with reference to FIGS. 5 and 6.

In one embodiment, the MTV has a height of approximately 20 inches and awidth of approximately 22 inches.

With reference to FIG. 8, the MTV 510 further includes a payloadinterface 550 that is located on a forward end of the MTV 510. Thepayload interface 550 comprises a coupling device or mechanism that isused to attach a payload to the MTV. In the illustrated embodiment, thepayload interface comprises an annular structure having a plurality ofattachment points that can be used to attach the MTV 510 to a payload,such as, for example, a satellite. FIG. 9 shows a schematic view of theMTV 510 attached to a payload 610. In the illustrated embodiment, thepayload 610 comprises a satellite. However, it should be appreciatedthat the other devices can be attached to the MTV 510.

There is now described yet another use for the oxidizer in a hybridpropulsion system. According to this use, the oxidizer from the storagetank(s) 215 in FIGS. 2 or 5 is used to drive a turbopump that is used topump the oxidizer into the combustion chamber 110. An outlet of thestorage tank 215 is coupled to the turbopump having a turbine that iscoupled to a turbine shaft. The shaft, when rotated, pressurizes a pumpthat is also coupled to an outlet of the storage tank 215. The pump canbe used to pump oxidizer from the storage tank 215 into the combustionchamber 110 of a rocket motor. Thus, the oxidizer (which is at leastpartially in a liquid state when in the tank 215) is routed out of thetank 215 and allowed to decompose into high temperature and highpressure gas state. The oxidizer is then expanded through a turbine soas to rotate the turbine, which in turn rotates the shaft coupledthereto to power the pump. The pump then pumps the oxidizer into thecombustion chamber. It should be appreciated that there have beendescribed two or more, three or more, and four or more uses for theoxidizer in a hybrid propulsion system.

A number of embodiments of the invention have been described.Nevertheless, it will be understood that various modifications may bemade without departing from the spirit and scope of the invention.Accordingly, other embodiments are within the scope of the followingclaims.

1. A propulsion system, comprising: a hybrid rocket motor case having aproximal end and a distal end, the motor case defining an interiorvolume including a combustion chamber at the proximal end, wherein themotor case is configured to hold solid rocket fuel; a fuel tank coupledto the proximal end of the motor case, the fuel tank defining aninternal volume configured to hold a fluid oxidizer; a partitioninterposed between the motor case and the fuel tank, wherein at leastone access passageway extends through the partition, the accesspassageway providing exterior access to the interior volume of the motorcase or the internal volume of the fuel tank while the hybrid rocketmotor is coupled to the fuel tank, wherein the at least one accesspassageway comprises a motor case access passageway, the motor caseaccess passageway having a first opening in communication with theinterior volume of the motor case and a second opening in an externalwall of the partition and wherein the first opening communicates withthe combustion chamber.
 2. The system of claim 1, wherein the at leastone access passageway comprises a fuel tank access passageway, the fueltank access passageway having a first opening in communication with theinternal volume of the fuel tank and a second opening in an externalwall of the partition.
 3. The system of claim 1, wherein the partitioncomprises a bulkhead flange sandwiched between corresponding flanges ofthe motor case and the fuel tank.
 4. The system of claim 3, wherein thebulkhead flange has an outwardly-facing sidewall and wherein the atleast one access passageway has at least one external opening in theoutwardly-facing sidewall of the bulkhead flange.
 5. The system of claim4, wherein the at least one access passageway has a plurality ofopenings in the outwardly facing sidewall.
 6. The system of claim 1,further comprising an injector configured to inject fluid oxidizer fromthe fuel tank into the motor case, wherein the injector is entirelypositioned within the fuel tank on a first side of the partition.
 7. Thesystem of claim 1, wherein at least one valve is positioned in theaccess passageway.
 8. The system of claim 1, wherein the accesspassageway includes a removable plug.
 9. A propulsion system,comprising: a hybrid rocket motor case having a proximal end and adistal end, the motor case defining an interior volume including acombustion chamber at the proximal end, wherein the motor case isconfigured to hold solid rocket fuel; a fuel tank coupled to theproximal end of the motor case, the fuel tank defining an internalvolume configured to hold a fluid oxidizer; a partition interposedbetween the motor case and the fuel tank, the partition including meansfor providing exterior access to the interior volume of the motor caseor the internal volume of the fuel tank while the hybrid rocket motor iscoupled to the fuel tank wherein the partition comprises a bulkheadflange sandwiched between corresponding flanges of the motor case andthe fuel tank and wherein at least one access passageway has a pluralityof openings in an outwardly facing sidewall.
 10. The system of claim 9,wherein the means for providing exterior access comprises at least oneaccess passageway that extends through the partition.
 11. The system ofclaim 9, further comprising an injector configured to inject fluidoxidizer from the fuel tank into the motor case, wherein the injector isentirely positioned within the fuel tank on a first side of thepartition.
 12. A propulsion system, comprising: a hybrid rocket motorcase having a proximal end and a distal end, the motor case defining aninterior volume including a combustion chamber at the proximal end,wherein the motor case is configured to hold solid rocket fuel; a fueltank coupled to the proximal end of the motor case, the fuel tankdefining an internal volume configured to hold a fluid oxidizer; apartition interposed between the motor case and the fuel tank, whereinat least one access passageway extends through the partition, the accesspassageway providing exterior access to the interior volume of the motorcase or the internal volume of the fuel tank while the hybrid rocketmotor is coupled to the fuel tank, wherein the partition comprises abulkhead flange sandwiched between corresponding flanges of the motorcase and the fuel tank and wherein the bulkhead flange has anoutwardly-facing sidewall and wherein the at least one access passagewayhas at least one external opening in the outwardly-facing sidewall ofthe bulkhead flange.
 13. A propulsion system, comprising: a hybridrocket motor case having a proximal end and a distal end, the motor casedefining an interior volume including a combustion chamber at theproximal end, wherein the motor case is configured to hold solid rocketfuel; a fuel tank coupled to the proximal end of the motor case, thefuel tank defining an internal volume configured to hold a fluidoxidizer; a partition interposed between the motor case and the fueltank, wherein at least one access passageway extends through thepartition, the access passageway providing exterior access to theinterior volume of the motor case or the internal volume of the fueltank while the hybrid rocket motor is coupled to the fuel tank, whereinat least one valve is positioned in the access passageway.